1{_ -.—

i’

DEVELOPMENT OF THERMAL ICE -PREVENTION

FOR THE B -24D AIRPLANE

By Alun R. Jones and Lewis A. Rodert

,“-;”,’,.

.NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

Wiurl’mll lmwr”ORIGINALLY ISSUED

February 1943 asAdvance ConfidentialReport

EQUIPMENT

Ames Aeronautical Laborato~Moffett Field, California

(,.

‘: ‘ti”ieA. ~~WASHINGTON

NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution ofadvance research results to an authorized group requiring them for the war effort. They were pre-viously held under a security status but are now unclassified. Some of these reports were not tech-nically edited. All have been reproduced without change in order to expedite general distribution.

A-35

-11 ,1- .m—m— ..- ... ! . ..—

31176013046235..—— __ _b ..-.

NATIONAL ADVISORY COMMITTEE.—

I’OR A18ROHAUTICS

ADVM90E COl!U1’ID~MTIALREPORT.,. . .. . .

.m!-DEVrniOPMEITT.OT $H.ERtiL IOrn.PEEVEITTIOITEQUIPMIllHT

~OE Tm B.241)kIBpLA”~ “. . .

By Alun “R. Jones and Lewl~ A. Eodert ‘

.,. .SUMMARY. . . .. . .

A thermal ice-prevention system for the B-24D air-plane has Ween developed” at the Ames Aeronautical Labora-tory of the Ndtic”nal Adri.sory Oommittee for Aeronautlosin cooperation with the Materiel Center. of the Army AirB’orces and the consolidated Aircraft Company. The sub:ectreport includes a description.of the design and an out-line of the method of dssign analysis. Results of per-formance tests of the installation are to be presented ina suppleiuentary repart.

The .thormal ice-prevention eys+tom is based upon rais-ing the tompe?ature of the surfaces to be ‘protected fromice formations by subjecting the inner face of the surfacoto a stream of heated air. Tho SUUTCOS of heated air arefour exhaust gas-air h~at oxchangors, one on eaoh engine.A doublo-skin typo of construction” WRS employod fcr thowings ~nd tRil eurf~coe, and dou?)ia-pano construction forthe windehi.olds. Tho hoatod air is Ceusod to Clrculatoby tho dynamic prossuro of tho air stream.

A design qnalysis is presented in a gonoral form, asa po8sib10 OUtlinO for futuro oomputatiotis, and is illus-trated with samplo onlculations from the B-24D airplanennnly~ism

IMTRODUOTIOE

In Cooperation ~i.th the Materiel Center of the ArmyAir I!’oroes, the Consolidated Aircraft Company, and sev-eral equipment manufacturing oompanles,. the Ames Aeronau-tical Laboratory has designed, installed; and tested in

flight thermal ioe-prevention equipment on the B-24il air-~lane . The work was undertaken at the request @f theMateriel Oenter “in order to relieve the aviation industry

.— ___ ___ .- —

.:-

. .

2.,0.:, :

of some of the tle8ign and development work whiob ia re-quired in the introduction and application of the thermalmethcd of ice prevention. It was desired that the equip-ment he developed so that produotien obuld” be undertaken “at once by the airplane manuf~otvrer. Attention, there- .fore, was given in the deeign to service, weight, produc-tion possibilities, and other such features. Mr. Howard .F. Schmidt, Consolidated Airoraft Company repre80ntatiVeat AAL for this project, contributed materially In the do-volopmeat .

Tho rapidity with which tho projoct was undortakonand complotod wae duo to the ~ntoregt and cooperation .given by all of tho intorosted agencies. . .

DESCRIPTION OX’T= ICE-PREVENTION EQUIPMENT

The B.:24D airplane is shown in figure 1. The airplaneIs a high-wing, tricyole-gear, heavy bomber pow~red by fourPratt & Whitney S3C4G engines, ratedat 1100 horsepower. Anexhaust-gas-driven ongino supercharger is locatod In eachnacollo.

Tho general layout” ●f the hoatod-air nntl-icing .Oystemdesigned for the B-24D ~i.rplane is shown in figure 2.Eented air is obtained from ?mexhsust gas-air heat ex-ch~nger in each nacelle. After passing through the ex-changer, the air is directed to the various regions to beheated by n system of thin-wall ducts. The dynamic pres-sure of the air stream augmented by the propeller providesthe source of energy for the cir,culatlon of the heated air.

The design of the thermal ice-prevention equipmentfor the win

!?outer”-panel leading edge (stations 335 to

626, fig. 2 is shown in figure 3. A“.spauwiee duct is .formed in the wing structure by placing a baffle at 4,5percent of the wing chord, and the heated air from theoutboard heat exchanger Is carried to this spanwise duetby the four-branch pipe system shown In figure 2. Thecorrugated inner skin and the outer skin form a series ofchordwise passages for the heated air. The air ontors thopassages through a gap in the corrugation at the wingleading edge and flowe along the top and bottom Inner sur-faces of the outer skin to the termination of the corru-gations at the front “spar (fig. 4). A series of reinforcedholee in tho front and rear spar webs allowo the air to

.— -

pass through the wing interior and out into the aileronslot region. The outor panel leading odgo, ready for in-stallation .on.th~ air.planer Ie shown in figure 5. 0

At the wing tip (fig. 6) “the air, after having paesodthrough tho Ioading-edgo eyqtam of the win% outor Panelsis allowed to pace into the forward portion of the tip andIH then mado .to flow hetwoon tho outor and -inner ekine.All of ‘the wing-tip hoatod air leavee the win% on the uP- .per taurfaoe, in. front of tho navigation light.

Tho thermal .Io;o-prwont.ion equipment design for thewing inboard-panel leading edge (etatlone 164 to 275) Ieshown In figure 7. A portion of the heated air at eachinboard exchanger ”outleb ie diverted to the inboard-panelleading edge, an shown in figure 2. The air enters atriangular-seotion spanwlee .duot Iooated at the front-sparlower flange, which runs the entire length of the inboardpanel . The air IB allowed to enter the chordwise paBeageO,formed between the suter ekln and an inner corrugated skinvhioh ie continuous around the leading edge, through aemall gap at the bottom of the triangular duet. The corru-gation paeaagea are eealed at their upper endO and the airpassee into the outside boundary layer through l/2-ineh-diameter holes in the outer skin. Circulation of the heat-ed air inside the wing is normally desirable, but was notfeaelble in tho oaee of tho Inboard panel because of thewheel-well out-out In the lower skin. The inboard-panelleading edge during inetnll~tion On the airplane IS shownin figure 8.

In ndditlon to supplying the inboard wing panele, theInboard heat exchangers aleo furnieh air for the empennagegroup and windehlelds. The. duct Hyetem in the wlnge andfueelage Ie ehown in figure 2.

The thermal ioe-~reventlon equipment design for theempennage group is shown in figure 9. The heated nir 18paseed through a 4-inoh-diameter tube mounted spnnwlse inthe etabillsor leadlng odgo, (see fig. 10. ) A Omall clotwae cut in the duet between etablliser ribs and tho thinduet wall bent Inward to form a sooop as ehown In tho slotdotnil of figure 9. A second ekln wne attmahod around thostabilizer leading odgo, extending 12 inohoa from the lead-ing odgo on tho top and bottom aurfaaea. Chordwimo epnoor8were employed to maintain a oonstant gap of about 0.051inoh botweon tho two lending-edge skins. All lighteningholee in the front spar wore so~lod with metal pl~toe. Aportion of tho hented air in the epanwieo supply duct paesos

— —

4

through; t~o stio”op”‘slot-s”‘i-n the ,dwot, “through ,holed”~iw thw , “l~tidi=~ “edg~ “uf th.b”l~?”er“skinr bo’twGGn. tho ~w~ “~ki~~ in ‘Rchordwi”se di%eo%iou, and ‘-over the upper and lower surf doesof the stabilizer behind. the front s~ar. The. quantity. ofheatd~ -air rema~nin-g” fn the Bupply duet is .~ischarged. ‘fromthe )duo.tiat t“he st”~bilizer tip .hnd pas.sd.sthrough holes ixt.the inb.o”ard skin of’“tlief~h Irito the fin. plenum ohamber. .

./...”-“ The fin ple”ntim ohambe’r,is “a”se~’led reg,ion, ”kor.me.dby ‘-”

plaoing baffles be”tween the webs of the fin ribs, and p-ro-.vides a praot ical method for passing the heated BIT fromthe stabilizer to the fin. The” plenum” chamber is shown.in figure 11. The “fin thermal i“ce-prevknt~on :h.’4~~n:t ,tieei~n is 8im21&r” *O that” of the’ gt~biliser.eki.n was.’tirhppe”daround the. leading edge to a diqtan~~ ~of9 inohe B f~em tlie.“leading edge, and spabers were empl.o~e.d.to prov%do. a conetan~ gap between skins of .0.0625 Inch,.Two 3-”d.n-o.hflexible ducts with outlets-at the .onds :only:, ‘wore fastened. to the plenum chamber to direct the air to “the’ top and bot~om ef. the fin. Li.ghtenihg holes ~n the :front spgr rind.the two end ribs (see fig. -9) were se”aleain o-rder .to retain all of the hated a-ir”i-n the l~oading-odge region and .foroe it through the” double-skin gap”. Tho .empennage-group, Yevteed for ‘thermal ice-prevbnt ien and ..Installed o.n the. airplnno, is” shown ~n f~~ro “12.

. . . .The thermnl ice-mrovention equipm~nt f“or the wind-’

shields Is shown. in figure 13; Protection $s”provided for.both the pilot’s hhd the copilot’s windshields, and the .heated-air-supply ducting is shown in figuro 2. .The wind-shield design oonsists of an inner plexlglas panel, read-ily removable In” Yllght”and spaced uniformly 1/8 ihch” fromthe” outer panel, and sn entrance nnd exit header .for theheated air.” The heated air flows spagwise a~rosb. thewindshields from the inboard edge outwardj and exhaustsfrom the exit header to the outside air stream throughslote .dut “into the forward edge of the side wind?w. . -..-

The”sxhaust gas.air”heat exchangers wero dhei~n;daround an existing portion of the exhaust-gas ta~l-stacklocat6d ’on the. bottom of each riacolle ‘between tho COZ1.OC-,tor ring and tho turbostzporchargbr. Sovoral hoat oxchang-ors of tho oxtendod surfaco ‘t~pe (having pine or fins prot-ruding into both tho:oxhau~t-gas and heated-alr rogi~n.s)havo boon tostod on tho a~rplano. Figure 16 shows Crio oftho typos tdstcd.

.,,. ,...’ . .. .~n”additlon to tho finned tube, the heat exqhaqge,r+ . “.. . .

6

r.”

consists of an aluminum Intake aboop (opening, 3 by 6in.~, a etqinless-steel shroud around the finned region,

.. - and-ad oUt3.e.t..manifold. The. q.ut.letof eaeh inboard heat‘ exhhanger. lnoludes”a dl&ht-angls’ bypass, with a butterfly~alvei f’or the purpoee of eupplying and controlling heated

“‘ “a”irto the wiag inboard panele. One of”the heat exchang-ers ita shown Installed In the airplane in figure 17P

Oontrol- of the thermal ioe-prevention taymtem was ac-complished by looating an sleetrio motor-operated dumpvalve in eaoh naoelle near the heat-exchanger outlet.(See fig.. ”S8~) The operating mechantsma for the inboarddump valvew ware emxtdnde’ilto tnolude the butterfly Valvesin the inboard-pan~l supply ducts. !Cho controls for eaoh“naoelle are Indepetideut; they kre located within reaoh ofthe oopllot and are ooqnected in euoh a manner that thedump valvee are always either fully open or fully olosed.The distribution of heated air in tho wing outor-panelsupply duot”s can bo varied by means of throo butterflyvalves in each wing,. looated in tho throo inboard (sta-tions 368, 450, and 515) honted-air supply duets near thefront spnr. These vnlvoa nro adjustable when the airplanais on the ground. The quantity of tair directed to theyindehiolds is controlled by a butterfly valve in the sin-gle supply line running forward in the fuselage. ThisVIIVO 1s normnlly in n fixed position, but it Is Rcoossi-hlo to orosv mombars. There aro no valves in the empennngehontod-nlr supply linc!s othor than tho inboard-nncollodump vnlvos.

DESIGN ANALYSIS FOR TH3!RUL ICE-PREVdZ?TION itQUIPMlilNT

A“deslgn analysis of the thermal ice-prevention aquip-mont was prepared to ostmblieh the dlmonsions of the heatedair passages and duets roquirod to produoe the desired tem-perature, air-flow distributions, nnd prossuro drops. Thegonornl proooduro followod In tho analysis is outlinod in

tho following pages, with the various stops numbered. Abrief discussion of tho pertlnont datn nssoaintod with eaohstop is prosentod, and actual oomputatione from the B.24Dalrplano onloulntions are prosontod as oxamplos. .

Tho fQllOWiILg notation was usQd In tho’qnalysls:. .

A aross-seetlonal s.rtm, square foot

8 surface Rroa, squaro foot

—— — — — .—. — -— .—

-.. .

6.,. . .,

. . . . .. . . . ● - . . . .

‘.. ..i.;-:. ..” -. -.. . 1

f ... . : “. . .;“ ..:

. . . ..: .“

“’:.%: .: tutal hbak”fioi~,.~tu per-hour””. ‘;”. ; “,: ..f .. . . .. . . .> .. . .,# . . ,.... . . . :“, .qn$t’hdat” fl~w, .?tu pez. hdur per uni:t length of spanq-.’””.’ - , . . . ..

‘ ‘ h “ I:.”-surface h~atltransfer oceffici,e?t,.“.. . Btu-per hour,a- ..

square feet, ‘r;..:. . ~“. ... ‘, “ . .. .

.- .-. .. . ~“.$emperatur?”~ ”””ti”. .t“ “

. . . . . ..T’ .iem*eriture”, ,03’ ~biilute . “: .’”” .. .

w;... .weight rate “of”:~lr f“low, poundspe’r hour ..,..a%.“. . . ... . .

.. ‘.w untt”welght rate of, air floti,”pounds. per:hour per unit.%tlength of” sp~n.. . “ .., . .... ..-. ‘.

.C specific “heat ef air, -.Btu per pound,P’ .‘F ..

. .

v“” absolute viscosity of air”, poundi pp’r second) “feet#

. .k thormfal conduc~ivlty of air, B’t.u per hour, squ~re., feet; ‘F-per f??t ;- “ .

G“. weight rate”of air fl& per.tinit of cress-sectional.-’nr.aa, pminds per second, ~qu-re feet..

P static pressure, pounds per square foot

;m

De .

.V.

“R

thicknese of” ghp between surfaces, feet

length of nir pRssag~ or””duat, feet,...

hydrnulic rR”diUs, or rntio of croes-eoct.lont=il nrento wetted perimeter in h duct,. feet

.... . .

equ~valebt diameter of a dkctl e“qu~l to 4m, feet.

speciftc volume, cubio”feet par .p.oun.d “ “ ‘

gas c“onstant (53.3 fo~.air) ..”.“ 2

acceleration: @f gravity., feet per se.con@&

f friction coefficient for air flow in ducts

s dis~anae as measured around wing leading edge, feet

v airplane indicated airspeed, miles per hour

c ohord, feet

I

—-

Sub~oript s:

8P refere to average iQn~itio”nO

z, a, 3, eto. are employed to define surfaoes or airspaoes and used as subsorlpts to indioate temper-ature differenoesl heat flow, and heat-transferooeffioiente. Thus the heat-transfer oceffioientbetween a given surfaoe (6) and adjaoent alr (“2)wculd bo written “ he-a s.. .

A fe”w ~ymbols ueod in the analye10 are not premented “In. the netation-beoause they do not appear throughout theoalculatione and beoauee their meaning i8 muoh olearer ifdefined at the place of’their. utae.

The analyale va~ based on an aasumed airplane indioat-ed airspeed of 160 miles per hbur at 180000-feet pressurealtitude .

SteD 1. A613ummt ian of fr ee-a ir temperature .- MostcaOes of” a3ro~~ft icing occur between the temperature of0° and 32° F. For the B-24D airplane analysis the valueof 0° F“was aaOurned.

&t~~. Aseumutlon of avera ~e temmeratu re llt which~b~ heat”ed surface Ie”ljo be ma intaine~ .- According to ref-erenoe 2“,paragraph D-6b, the temperature rise ●ver theforward 25 percent of t4e wing chord must be at least 70° Yabove ambient air, and the rise between 25 and 75 percentof the oherd must” bo at least 20° F nbove ambient air. Di-rect heating of the leading edgo for 25 percent of theohord is difficult to obtain in oertaln designs, and insuch caees.tho resumption Is”made that by heating directlya smaller portion of the loading edge (say 10 to 15 ~or-cant) to a tempor~tura rlso of 100° F (instead of 70 F)nnd then discharging the hoatod air to tho remainder oftho wing,” tho spocifioations of reference 2 can be snti~-fied. The Lockheed 12A “of reference 1 is nn example ofthis compromise In design whioh hae proved capable ofproviding ioe prevention. Direct heating was providedfor the forward 12 percent of the wing, raising the wingtemperature hpprcximately 100° Y, and the heated air WRSdischarged from the leading-edge region and circulatedthrough the remainder of the wing.

Direot heating of the B-24D airplane wing was limitedto the forward 10 percent boonuse the loo~tlon of the front

8

spnr at that pofnt prohibited further extension of thecorrugated inner ekln. The corrugated region of the skinwee sub~eot to thermal analytic~l treatment. The hented-wing design wns based upon a 100° B temperature rise everthe forward 10-percent-chord region and an Indeterminatetemperature rise over the remainder of the win-g. A tem-perature rise o.f the wing nfter-~ortlon will ooour becauseof tho bent.ed boundary layer, nnd the digoharge air. fromthe len~ing-ndgo ByBtQm. In the ease of. the wing outorp~nel, thp hentod air was disohnrged through the” front sparPnd c~rculated ip the wing interiors .simlln”r to the Lock-heed 12A nlrplane design. For the inboard panel, however,the heated air could not be circulated In the wing lnterl-cr. The heated efr ,was therefore releneed to the boundarylayer on the upper eurface ahd carried back over the wing.The stabilizer and the fin were treated in a manuer simi-lar to the wing inboa”rd panel, with the exception that theheated air wae diaoharged ove”r both eurfacea of the air-foil sections. The doublo-ekln eystem for the empennagewas dictated by the simplicity of the revisions requireden the existing empennage, and the desirability of produc-ing a auction at the heated-air exit to aid the heated-~ir flow. The horizontal stabilizer loading-edge akintemperature waa ra.ifled90° F, and the fin leading-edgoskin temperature wsa raiaed 70 ‘“F in the design; This .heating ef the ampbnnage surfacea should bo adequate bo-cauae the discharged air frcm the leading-edgo.system iaoffoctively diatributod ovor both aides of the airfoil.

Steu 3. CR~culation of tho hea t-tranefer coefficient.&otwoon tho wing surfncc an d the.a mbiept air.- With thotompornturo difference bctwcon the wing akin and tho ambi-ent ,air dstabliahod, tho quantity of hoat romovod from thowing ekin dopcnds up”on tho outor-surface heat-tranafosooofficiont. Tho coefficient can ho” calculated from thodntn in roforonoo 3, although this method involves aomoerror bocaueo of tho low Reynolds number at whioh the tostaworo made. ~nothor method of dotormining tho ho~t-traneforcoofficiont, based upon tha rolntird botwoon hont transferand vigoourn.drag, is presented in reference “4. For theB-24D airplane nnalyaia, the data in”refsrence 3 were em-ployed by extrapolating in the equation

(1)

where the first prime refers to valuea from referenoe 3,the second prime to valuea for the B-24D .qnalyaia, and n

7.

,.

Q 6-?= (h

8-7)BV ‘(te-7)~v “ (2).. . .

(See equation” (1) , p. 136, referenoe 5.) Form the B-24Dairplane-wing outer panel, the design surface was the en-

“. ti~e’ ‘outet-pan~l “stirfac&for.wnr’d”cif the front ~par. !l?hediwtamaae wvoand”..tha tilhg leading edge from top to bottom ,

‘of tho front. ~par” is shown. in figure .19,.and the average‘value.for-”.th”e.outsr;-pamel mul.ti~lled- by. the panel span ro-

“’”faUltadin:ti OnrZnno “nr.aa:“of 43;9 ‘aquttr.u:foot.i .~rom fig-“uro lSIT‘l”h’~~.-for thei~qta;.~~~el: 3.,13.Btu.poa hour, equaro

foot , ‘r. Then .i::~.... .{.,* ~1- “, . .

,. . .-.

.,-. ‘:St:ou:‘5.”-iuOti‘~’atiob ‘h”f:‘t-h!o‘a;m”bUnt’”of h:e’Rtthkt “uhouldO.-vla a .i=i’i~p%lbin th~ alr

—to

..etzr0“the .noce”~gar~

ho”a.tdaqw to tbo e?ur~a ~~- ThQ:.mqu~ri#:ityof h-oat.thnt must... be’ eUppl~Q& ta.%h’e”“dea~gn sti,fa~e.,d~ ~&cen ,t.ab“e from two

to four t<mcte-fiha he.~t flow fr.04 ttit eurfacs”,.de,ppnd~ngupon the amount of heating the air is expeoted to deliverafter leaving the leading-edge region. In the B-24D outer

10

,

.. .“ -. , ., .: :... .:

~.. . . panel+tlie to.tbl heat to be eupplled to the lea.d.~ng-edge

rbgibn. w“as.assumed to be 200.000 Etu per hour”. r “-.,. .,..... .‘.. ‘, . . .,* “Stem 6. . AomumDtion of a temmeiature r“lie “for the

ip Daf@w thr~. .

Ltinted air.. the heat -eZoIa@zer..- Th.k de--.,

~ign temperature:. rise through the he-at.,e.k.chqn”geri:s..usually. ,de”tjermlned by the..maxirnum temperature allowable for the“heat. excha~gek and heated-air-duct materials+ and poskible. . .effects’ of e.lp+a.ted temperat.wee on th~. a~rglene primary

..etrtieture. A. h.e~ted-atr. temperature OY 30? F at. tihe.ex-chang.o”r’”putlet.yas, considered, ,to %e a .reas.enable and. safe

“ ilasi’gnvalue for the B-24D airplane. . , . ,

. ., ,. ,. .still 7. Oal~ation the. rate it of:airof weie floy..

r.oaulre& .- The quantity of air that mugt pa&s through” theheat exchanger in a given time, with a temperature risoestablished by step 6 in. order..to prfiduce the availa31e

“ heating of stop 5, is doterrnined by.,. ..

., Q.wE— (3),. Cp dt. . .

whore At is the heated-air temporaturo riso in ‘F. B’ortho B-24D outboard hoat exchangers using tho valuo offrom figuro 21, . . CP

w= 200000= 2730 pounds por.hour

: 0,24 X 300

Stou 8. Dosign of the heated nir na~ ●- The de-sign procedure employed was to divide.the wing surfaceinto unit strips running chordwise and cons$der the heatflew for a suffiblent number of stripe to define the wing .heating. The width of the division strips for the outerwing panel was taken as 1 inch, or the. Apahwise distancecovered by a single corrugatltan, .

~.

Sten 9. Anpli&ion of Stel)s 2. 3,-d 4 to ind IvI&-

Ml ch~ise stri RR.,- The heat flow from the outer-sur-face nren of each strip considered is determined by appli-cation of steps 2, 3W and 4 1P the same manner a.s employedto determine the total heat flow f~cm the total criticalsurface . The single corrugation’.strip at.the inboard edgeof the outer wing panel (station 335) will serve asian ex-ample. The heat flow from .the outer “surfaceP .

)’

.,

!LS-7 = h13-7 te.v &

h 6-7 = 11.6 Et; per hour, square feet, ‘E

e = 2.46 feet

ta-7 = 100° r

., :“

11

(4)

(fig. 19)

(fig. 19)

(step 2) ..

11.6”;X io.o.x;*”=.- .-

,qd+ = 237 Btu per””ho.un” ~....“ :.... ‘-“ . . .

.. ... ,:. , -.. . . .... . ,.- “... . . . . 2 .

@Q “l@:” ~~~~mmti~n o’f yei~ht di~tribu{ioti if ‘heatet

ti*- ?he. dis~ilbut:on of-thd.heated air to the various airpassages Iei determined by trial and error, considerationbeing given to suoh faotors as the larger heating requiro-mont at the wing root, tho inoroased surface transfer coef-ficient nt thq win.g”ti~, and the preesuro drop in the airducts ~nd paeeagen. The final weight distribution whichprovod ~t?tiafactory in the airplane anb.lyala wna to supplythe air .to the corrugntiona in. quantiltlea Invoraoly propor-tional to the square root of the distance oround the lend-ing edge, or

w = WRTf

%Ea

(5)

Tho B-24D nirplnno-wing outer pnnel has 291 corruga-tions; thorofore tho value of w nt station 335,

w= 2730

1

~= 4.1 pounds por hour2 X 291 2.45

Ste~ llm- Calo ulation of the temue rature droD of the

~~$ed air in the chord ;iee uaeaauea .- The quantity of airflow~ng in each pasO?ige” and the heat removed from the airhaving been estnbllsbed,. the temperature drop of the heat-ed air can be calculated -by applying an adaptation of equa-tion (3). For the oorru”gatlon at atatlon 335,

q e-7237

“ tl - t3 = = lle” r2cpw = 2 X 0.24 X 4.1

6ten 12. Dea3~n of the heate d air D~e to DrOdUCethe reo ulred heat floq .- The remaining step ie to vary the

. . .. . . ....

.,. .-

12 .

Reynolds number of the heated air by ohnng~ng the ai”r-pas-sage dimensions until a heat-transfer coefficient. for th~in~ide of the passage is produced which will supply theneceegary heat to the outer skin, with the average tanpera-ture of the air in the passage determined from step 11.The Eeynolds number for the flew of the heated air is ex-pressed: . . ..

GDe.Ri = — (6ee pp. 99 and 2350 reference 5.)... w.

. ...The heat-transfer coefficient for the inner surfaca “

of ‘the Fassage i~ determined from empirical data showingtha-varlati.on.of the Nu~self number with Reynolds ”numbar .whero the NusseZt numbar

. ..“

~~”. * (See p. 96, ref”arence 5.) . -.

~mplric~l data which havo proven satisfactory in”de-termining the air-passage heat-transfer coefficient aropreecntod in figure 65, refo~oncb 5, qnd tho recommendedcurvo AA from th~ flgur~ has been reproduced In figuro 22.T.hc dat~ dot~tinining tho.curvo AA in figure 2.2 nro diroci-ly concornod With fluid flow In circular ducts, but expe-rience hes shown that reasonably accurnte calculations”for ducts of noncircular croes section can be based uponthe curve AA, provided the departure from a circular crosssection ie not too severe. In designs where tha Rig pas-sfige consists of two Parallel plates, such as the empennagemnd windshield design for bha present ai-r.plane, the empir-ical data presented in figure 7 of reference 6 and repro-duced In figura 22 are recommended. Thesa data nro plot-ted on the bssie of the gap width d ~e the aquivalant

. dinmeter De, and their compfirisen ~~ figure 22 with therocommanded. curve AA from refer~nc.e 5 reveals the errorthat can be introduced by applylng curve AA to air pas-sages of nencircular section.

In tho caso of tho B-24D airplano wings, tho hoat-transfor coofficiont insido tho corrugation air passagcewas bnscd upon curvo AA, figuro 22, Consldoring tho singlocorrugation at stfition 335,

G=~ 4.1 = 1.49 pounds por second,3600 A = 3tj0CIX 0.000765 equnro faot

—. .—

of M from $Iguro 21,-. .

. . .. . . . . . . ,.

From cutivo:&A, figuro 22, ?or Ro ~ 1710,...

ha-:e”Do = 7,6..HU =. .. . ... . . ..k”

..

From fi.~ro 2?, k = 0.0162, “and thoroforo, .“.. .

,.,.

. . .

. . .. . .. .

-.

7.6 X 0.0162h“=— = 7.3 Btu por hour,” .a-0 4 x 0:0043 Squnro foot, ‘r.,

Tho avorngo tomp~raturo of tho sir In roglon 2, assumingt~ = 3QQ0 r,,

“t=118

=300-y = 241° P’Rv

“ The qunntity of he~t flow to tho skin then bocomos

qa-a = ‘a-e x + x ‘taav - te)

= 7.3 X y x 141

= 210 Btu per hour

This value of qa-e is In satisfactory agreement with the

value of qe-7 equal to 237 Btu per hour determined In

step 9. The thermal design for the wing inboard panel andthe empennage was made in a manner similar to that usedfor the outboard-panel design, and the resulte are shownin fi.guren 20 and 23.

In the oaee of the B-24D airplane windshield, effic-ient data were hot availabld to determine the heat-transfer oooffibient from the outer surface. Steps.2, 3,and 4 of the deeign procedure were replaced by the aesump-tlon that a heat flow from the outer surface of 1000 Btu

14. .

.. . .

per hour, square feet would be sufficient to provide ice -“Oprotection. (See p. 8, referenoe 1.) Further neeossaryassumptions were the temperature ef the windshield outerpanel, ~ssumed to be 50° 3’, and the temperature of the airentering the gap between the panels, assumed to be 150° S?.The design wae then completed by assuming different quan-tities of air flow and values of gap size until a combina-tion was found which would produce the required outer-panel heating. The results of the windshield analysis areshown in figure 20. The desired temperature rise for thecritical surfaces having been established, the pressuredrsp in the air paseageta nnd the deBign of the heated air-Bupply ducts to obtain the necessary weight of flow distri-.bution are exRmined. Unless an air pump of eeme sort isincorporated In the thermal anti-icing system, the circu-lation of the heated air is dependent upon the total en-ergy of the nir at the hep.t-exchanger inlet. Seaondaryfactors whiah mny be considered to aid the propulsion ofthe nir through the Bystem nre the ~ddition of energy tothe nir in paseing through the heat exchanger nnd the lo-cation of the air outlet at n point of low pressure. Fortho B-24D thorml Ice-provontlon system, tho energy offoctIn the heat oxchangar was neglecto~ nnd a prossuro drop of5 inches of water, or approximntoly ono-half tho valuo oftho dynamic pro~suro for tho doslgn indicated ~pood of 150mllcs per hcur, was n.ssignod to the hont exchnngors. Inordor to obtain tho weight distribution of air dosirod intho surfnce-honting cnlculntione, tho prossuro drop mustbo oqunl ~long all hoatod rir paths, from tho hont-oxchangor outlet to tho nir exit from tho wing. Tho gon-oral design pc~coduro is to onlculato tho pressuro droproquirod in tho various air passages nnd then design thosupply ducts to produco equal prossuro drop in nll of thopossible nir pnths. .

Tho prossuro drop in the he~ted air passnges and thesupply ducts was calculated by the eau~tlon

(6)

(equntion (35) , P. 1~0, reference 5), where the subscriptsz and a represent the extent of region over which theressure drop is c~lculnted,

?)The first’ term @f equation

6 was found to be negligible in thih case snd there- “fore dropped from the equation. Data from whioh the fric-tion f~ator fav mny be obtained nre presented in figure

—

16

24, -whiali 1s n reproduction of figure 8, referenoe 6. The

. .c.urv.d‘~.nfigure 24 refers to .-flow in oiraular pipes, andthe .dntn plotted Sn the figure were obtained i&.modeltests with air flow in R narrow gap between p~r~llel sur-faOOBe Figure 24 w~a employed in the analyeie”in.tho aambmmnor “nEI- figure 22; thnt ie,.tho ourve for Oircmlar .pipoewaeFused when determining. f for thet oorrugwtod air p~m-eagos and 6UPF1Y duets, and the plottod. dnta woro keod ‘when determining f for”nlr flow In tho ompennngo andwind6hiold gaps: . . .“

OonBidoring ~ einglo Corrugation a~r”p$tee+o Rt Otn--tion 335, outer wing panel, ., .“

RT 53”.3 x 701 x 2‘av = ~ = w 35.4 cubio feet per ~otind

2116

The static preesure at 18,000 feet wae used in oaleu-lating the specific!”volume. ‘

Re =G De 1.49 x 4 x 0.0043 x 10= “ = ~vlo “ “—.a-

f aV = 0.0094

n g=‘2

1.22

w(3—= 3600 A =

1.5 .— —

(from curve In fig. 24)

feet

4.13600 X 0.000765

= 1.49 pounds per seoond,square feet

m = 0.0043 feet

PresE&e drop from region 1 to region 3,

f mla v ~P, - P= =

0.0094 x 1.22 * 1.49a x 35.4v2gm”= 2 X 32.2 X 0.0043

. .P’z L P3 =“3-~3 pounds per or 0.6 inoh-effiwaterq.

square foot

Extension” of the pressure-drop oalbulations to other cor-rugation air passages in the wing outer panel indicatedthat the preesure drop from region 1 to region 3 was sub-stantially constant for the entire panel. (see mea. u,

—. —.

16

fig. 20.) In order to produce the.desired distribution ofthe heated air, therefore, the pressure drop from the out-board heat-exchanger outlet to any point in region 1 hadto be constant. The spanwise pressure drop along region 1was found to be too large if all of the heated air wereadmitted to the region at station 335, hence the air wassupplied In four tubes as shown in figure 2, .The pressuredrop through the corrugation passages ‘in the inboard-panelleading edge was caleulnted to be 5 inches of water asshown In section BB, figure 20. This large drop WAS con-sidered allowable because of the looation of the heated-air-exit holes in a low-pressure region. B’or the empen-nage group (fig. 23) the fin heated-air gam wae designedlarGer than the etabllizer gap in order to approximatelycompensate for the mreseure drop in the stabilizer leading-ed~e SU??lY duct.

Instrumentation of the B-24D Airnlane for Tests

Thermocouple, preesure orifices, and venturi meterewere included in the design of a portion of the B-24Dairplane thermal Ice-prevention equipment in order tomeasuret8StSmterest:

1.

2.

the performance of the installation in flightThe following factore were considered to be of ln-

Quantlty of air flow through the heat exchang-ers and various parts of the equipment,

Temperature of the heated air throughout the eys-tem.

Temperatures of heated surfaces, namely, wing andempennage outer surfaces, parts of the Internalstructure, windshield panels, and heat-exchangersurfaces,

Temperature of the exhaust gas.

Static and total pressure at the heat-exchangerinlets, and static preesure of the heated airthroughout the syetem.

Exhaust-gas static-pressure drop through the heatexchanger.

The quantity of air flow was determined by the use of

—..-— —

.:- . . .

. .

., 17. ,. . ..-. . . .. . . . . .

venturi meters, Four “such meters were insta31ed, ae. ~hown,, in figure 2: cone in .+ha S-i.noh supply line”fzom the right

outboard ,hent’exchanger, one in thb Z-Inch inboard-panelsupply lin’e from” ths right inbokr~. heat Bxchangerj, one Inthe 2-inch supply duct” to the .copilotle windshield,”.andone in the 6-inch supply duet to the em”p.ennage (locatedaft qf the Junction of the ducts”from the ‘Imboard heat ex: .“ohangers) . The ratib.of the throat “diameter to pipe diam-eter for the venturi meters was O*7O

All temperature readings.weqe obtained with iron-oonstantar+ thermocouples and a Lewis potentiometer. Theidentlficat.lon drawing tor the thermocouple Is shown infigure 25. The dash numbers following the thermocouplenumbers in figure 25 refer to .thb type of thermocouplemounting, as detailed in figure 26. The Plexiglas shieldfor type 5, figure 26, was required beoautae “the thqrmo- . .

couple junction” otherwise would move away”from the outerskin and protrude Into the ambient air stream..” Two thermo-couples were located in the intake scoop of the rightInboard hea”t exchanger,

The locations of the pressure-measurement points areshown in figurd 27. All meaduremqnts were of statfcpressures, with.the one exseption”of-the total pressure Inthe intake scodp of t4q righ$ inboard heat exchanger.Three total-pressure heads and %WO” static-pressure headswere distrtlmted aoross the exchangeT inlet because pre-liminary flight tests. revealed a variation In total headIn that region at low angles of attack. !I!hedash numbersfollowing the pressure-orifice numbers in .flgure 21 referto the type of orlflce mounting asshown in ftgure 2g.

All pressures, with the exception of the exhaust-gaspressures, were referred to the total pressure from thepitot-statio airspeed heads located at the nose of theair~.lane. !Che pressure differentials were indicated bywater manometers and airspeed indicators. The abeolutevalues of the two exhaust-gas preesures were indicated bya maalfold pressure gage. A calibration of the differencebetween the static “pressure at tbe airplane alrepeed beadsand free-stream static preesure was obtained by suspendinga trailing static head from the airplane, The static cal-ibration of the airepeed heads provided a basic for refer-ring the test pressures to free-stream static pressureand for determining the correct Indioated airspeed of theairplanea

I

. . . ...-

. . - ““. .

. .., . .

... .. a. .

lg .,. . . .:.”

..

.: .

..?~~LIMI~J@y ~LIQHT TESTS - RESiLTS AND DISCUSSION1. ..

. .

Prelimlnary,$llghts have been conducted with the B-24Dairnlane’to” teat the performance of the thermal ice-preventi”on equipment. Supmlylng a qu~ntity of heat to the .wing outer panel. equal to approximately 65 percent of the . .design quantity produced a temperature rise of the skin for-ward of the front spar elightly in excess of the design . .value. Indications are that the remainder of the thermal “ice-prevention design will be equally satisfactory when he~tquantities approximating the design values are supnlied fromall four” heat exchangers. Data on the performance tests ofthe complete installation will be presented as a summlemezt=-tary report later.

Weight of Equipment. .

Calculations have been made to estimate the increasein weight ofia 3-24 airplane resulting from the instnlla- ‘“tion of a production modification of the subject thermal~ce-prevention system. A study of a“productio~.designwas considered more deeirable than a presentation of thewei{;hts of th6 B-2JD airplane installation because thati.nstallat~oi ie m experimental revision to an existingairplane and the weight faator was not given the. consider-ation that it would receive In a production design. Thecalculatloms indicated that the weight of a B-24 airplnne(not equipped for Ice protection) would be increased “about300 pounds by the installation of-thermal Ice-preventionequipment. Attention Is called to the fact that the figureof 300 pounds “is subject to revision on the basis of main-tenance and durability requirements as determined by themanufacturers” experience, The JOO-pound weigl@ of the .present equiqment oompares with the 2J0-mound weight ofthe inflatable de-~cer equipment that it remlaces”. Thelatter does not include windshield de-icing, however.

..COMCLUSIOMS

1. Y!hermal ice-prevention equipment for the B-2JDairplane wings, empennage, and windshield 10 structurallyfeasible.

2. The thermal ice-prevention eouipment installationwill probably satisfy all design requirements-

19

3. A modified produotlon installation of tho thorzmlIco--prevantion eystom would Inoreaso tho wolght of R B-24.airplane (not oquippod for ioo protootion) about 300 pounds.

AnotI Aoronautlo~l Laboratory.National Advisory 00nn~ttao for Aoront’tutioe,

Moffott Field, Cnlif.

REI’ER3!HCES

1.

‘3,8 ,

3.

4.

5.

6.

Rodcrt, Lewis A,, Clouaing, Lawrence A., and McAvoy,William H.: Recent Flight Research on Ioe Preven-tion. IITACA A,R.R,, January 1942.

Army Air E’oroee Specification No. R-40395: Anti.ioingEquipment for Aircraft, General Speoiflcations(Heated Surface ‘Yype). A.A.F,, April 21, 194?.

Theodorsen, Theodore, and Clay, William C,: Ioe Pre-vention on Aircraft by Keans of Engine MxhaustHeat and a Technical Study of Heat Transmissionfrom n Clark Y Airfoil. Rep. No. 405, NACA, 1931.

Allen, H. Julian, and Look, Bonne C.: A Method forCalculating peat TrRnsfer in the Laminar FlowRegion of Bodies. NACA R.B,, Dec. 1942.

McAdams, William H.: Heat Transmission. McGraw-HillBook Co,, Inc., 1933.

Rodert , Lewis A., and Jackeon, Richard: PreliminaryInvestigation and Design of an Air-Heated Wing forLookheed 12A Airplane. KACA L.R.R., May 1942.

““%’

F

Figure l.- The B-24D airplane in which thermal ice-prevention equipment on the wings, ?empennage, and windshields have been installed. P

I?ACA Fige 2

~

In..,:;

-.,------——

I __________ —

-’””w+-.. \

.-. ...‘\

‘w ~-- ..

)’..,.,

\’. ~~~

3“.-.....\+.

*’

-’%

“u

NACA

-.

Figure 5.-The rightwing outerpanel lead-ing edge .for the B-24Dairplaneafter thealterationswerecompletedfor the in-stallationof thethermal ice-preventionequipment.The photo-graph showsthe heated-air supplyinlet holesthrough thebaffle plate.

Figs. 4,5

Figure 4.-Details ofthe wingouter panelleadingedge of theB-24D air-plane duringalterationfor thethermal ice-preventionequipment.Shown arethe air gapat the leai-ing edge,nose ribconstruction,and the bafflesupport angles.

fOWER saw

=%?+. .

~,u.vfin 5M!N (Ewmvd 4.

(.- 1’--L’

—2

~.. .--l

1 \\

rF#ONT S48.4

@/K-7 /.

.4-.71s r/.vG

lfOL Es(TOP eBO 7r0m)

—

.?/v.,4u7s

.5uR&Ac4a7sTIN6

RIB /VT sm.”PU 1 1

SCAU /#”. /:@.

6%SEC TION ? -C

SOLE !2-.l.b -

.4aw Ouma Ww (04a ,s=’.+”

II‘,u~ ii I

+

,,

‘i 1

---------- ●----- M- Li~—-–—--—-L ----–––-ET

---17 ---- 7--------- ?l~---------, ------- -

,,

,..

::,1 I

!Jl ,1LQ_J_l__ .— .m+_%,_._LIll

PLA?J “VIEW OF LEFT WING TIP5/4LEa SCALC 3.. l+Y-

/-#uLKUf4

-

—

E

~ wrm sam @fO”.3sd*l

Figure 6.- The heat-ed Wing

tips of the B-Z4Dq

airplane showingthe double skin G.leading edge details. ~

*

SECTION .4-ASC.4E 12 ~ 1:0- SC4E 12-4:0-

., :!.

\

,7

. ●.

.

>“)

\ I -n-.Figure‘7.- The Wing Inboardpanel leading edge

design for the thermal Ice-preventIonequip- /~“

dment on the B-24D airplane.

NACA Figs. 8,10

Figure 8 .- The left wing inboard ~anel leading edge of theB-24D airpl~e shown during the in~talxation of

the thermal ice-prevention equipment.

Figure 10.- The horizontal stabilizer of the B-24D airplaneduring the installation of thermal ice-preven-

tion equipment, showing the lightening holes in the nose ribsthrough which the heated-air duct was installed. The stabil-izer is viewed from above in this photograph.

,Z/G?f TStfING HOLES

/

/’COVEKEOIW i_HIs .71B ANO

.71 MILAR ma ar~orrcw —~

Y “’ “ 1

.

Figure 9.- The therma[ ice-

pravention equipment

for tne empennaqe group on +he E21

B- 240 d-plfme.

-

L+! -m. *T 4.4 #uC,-ON zJs.#w4u#

PATH OF HEATED AIRIS sh#wN BY ARROWS

~~

.—m-— !!.- .-1. I . . . . . . . -- . . . I , ,, . . II .!-. ..!. -.. ! ., .,,... . . . . . . . . , . . . . ,,. , ,,. . . . . . . . -,-- .—. -.-...— ——-.——

Figure

thermaland air

Figure

Figs. 11,12

11. - The inboard side of the right vertical fin of theB-24D airplane during the Installation of the

ice-prevention equipment, showing the plenum chamberducts running to the top and bottom of the fin.

12.- The empennage group of the B-24D airplane inwhich provision has been made for thermal ice

prevention.

Figure 13.-The double pane air-heated ~windshield installation for ~

the pilot’s’and copilots’ windshieldsthe B-24D airplane; Thereadily removable bv.the

inner panelpilots in

IIT’_fl====+-%=&L_J-...

“’TW?NSMJllh ‘

----- _ --

of “is

NACA Figs. 14,15

Figure 14.- The pilotls air-heated windshield on the B-24Dairplane, showing the inner panel partially

removed.

Figure 15.- The pilot’s air-heated windshield on the B-24Dairplane, with the inner panel secured in

position.

—

Figs. 16,17,18

Fiew?e 17.- The exhaust-air heat ex-changer installation on

the B-24D airplane in the right out-board nacelle.

Figure 16.- A section of the B-24D airplane exhaust gas tail-stack which has been converted to a fin-type surface

heat exchanger for use with the thermal ice-prevention equipment.

Figure 18.-The heated–airdump valve forthe right out-board exchangeron the B-24Dairplane. Theposition of theelectric mbtor-operated valvedetermineswhether or notthe heated airenters thethermal ice-preventionequipment.

.

\

$-l

*

m.

30 160 240 320 400 480 560Semi-span,in.

Figure19.- Curves showing the distancearound the leading edge from topto bottom offront spar, and the heat transfercoefficient,

I

I

I

for theleadingedgeof thewing surfaceon theB-24.Dairplane.

Pco

m AMxzw- )AWG.L=

Figure 20. - The results of 7the analysis

of’the thermal ice-prevention equipment} showingthe distribution of theheat,edair in the wingsand windshield in theB-24D airplane.

.252 5.OX1O-5 3.2x10-21

.250 4.0

CL .

6-.246

.244

.242

32.0

1.0

0

4-— /“

— /

// /

,/-.-

//

/

=>~

1.20 200 400 600 800 1000 1330

Tigure 21.- The p’hysicalpropertiesof air as employedin theanalysisof the thermalice- %K.~

!i’ernperature, ‘F

preventionequipment.

7-

/

+

+-

--1-=1=—

k

——

,--

.——

“=4

.—

—.. -

/“ !

. .

- --U -41-------i----%FFFFF:7+--

2 3 4 56 8 10 2 3 456810 2..”

Re = 9, consistentunitsw

NM

NACA I?ig.23

7

6XP =

/ r-A \\ / \

c---__._/___ l__/___ ‘__+ ___/_ i____. ____/_______

T—-———-—-————---.—-- --- —--____--—-—-—-—

/365 JL@%ff “OE#/? AT

%%%WL%if’’%#/%R -A

,

)t-”-i

Figure 2%.- The results of the analysis of the thermal ice-preventionequipment, showing the distribution of the heated air in

the empennage in the B-24D airplane.

‘-q.100

.080

.060

.050

,040

q+ .030

*“$.rlo .020c‘l-l:c)

~%-1-P.2# .010

.008

,006

.005

.004

Re =DeG—, consistent units

P

Fi7gure 24.- The relation between the friction factor and Reynolds nrurber for flow of air in pipes,(De = diameter, ft): and in narrow gaps between parallel surfaces, (De = gap width, ft)

Taken from figure 8, reference 5.

4

gi..’

NACA

-%A 33-4A.3/ -5 /A 32A 30

/V6

A Z9

zP75-7 —4,?8-5A27-A26-

.444-.4

“IT”+ p’+

“1; ——-n-*’ ~@$*.._....fT7==”.=”-- ---~.-..;,%

i!%!k+ ~ >..—-.——_-s ‘~,~.r33-2

#7-5

-/ + J34 -2J-3S -2

)-? Jo”/ - f

/=//’v

Figure 25. - Theandthe

thermocouples which were installed on the B-24D airplanewhich were employed in studying the thermal qualities ofice-prevention equipment.

s. ..,,!.._

rANUS 403-4

● m-rS4FC3-)ON2-4’

‘“4::“’e:’’’’””’+L 7am?.4T0cuAE Ju~r~~

~

26.- Thermocouple installationdetailsvention equipment.

=... 0%.. I 1i u 1 w

\ TZ?Z40COUAE .

TYPE J]

in the B-241)

Fig. 27.,

—

Figure 27. - The pressure orifioes which were installed on the 13-24D air-plane and which were employed in studying the thermal quali-

ties of the ice-prevention equipment.

Ii —

. r,. .

--Fl~re 28. - Pressure orlf Ice detal19 employed

mZ!s 6 in the B-24D a!rplanethermal 1cc-preventionequipment.

4“.4?-.7 \ / /-*,cn3

*Y %.p’ml. 1